Cooled turbine rotor blade

ABSTRACT

A cooled turbine rotor blade for a gas turbine engine includes an airfoil and a profiled root radially inward of the airfoil, a platform between the airfoil and the root, and a cooling air supply channel extending through the platform into an interior of the airfoil and therein up to an outlet opening. An inlet opening of the cooling air supply channel is located at the rear side of the rotor blade, and an inlet part of the cooling air supply channel with the inlet opening is angled into the direction of rotation of the rotor blade and curved into a radially outward direction. Further, a rotor-stator stage for a gas turbine includes a rotor blade as above, and an air cavity radially inwards of sealing features between the rotor stage and a neighboring downstream stator stage to form a source of cooling air for the rotor blade.

This application claims priority to German Patent ApplicationDE10208130897.3 filed Dec. 4, 2018, the entirety of which isincorporated by reference herein.

The invention relates to a cooled turbine rotor blade for a gas turbineengine, in particular an aircraft turbine engine, and to a rotor-statorstage for a gas turbine providing such a rotor blade as described in theclaims.

In modern aircraft engines, it is standard practice to cool the turbinerotor blades that are exposed to high temperature. Turbine rotors withcooled turbine rotor blades are known with different designs of coolingair supply channels, wherein it is referred to EP 1 004 748 B1, EP 1 464792 B1, and EP 2 551 458 A2 for example.

All these solutions have in common that they aim at providing amaximally effective cooling of the turbine blades with a high-pressurecooling air that is supplied to the turbine rotor in order to minimizethe temperature of the turbine blades and to ensure a maximally longservice life. As the cited examples show, it is typical that the area ofthe turbine blade intake at the disk rim of the rotor disk of a turbinerotor is designed in such a manner that a slit-like cooling air supplychannel remains between a blade root of a turbine blade and a groovepresent between the disk fingers of the rotor disk, which support theturbine blade. A cooling air flow is guided through the slit-likecooling air supply channel in the axial direction of the turbine bladeand its root portion, and enters cooling channels provided in a aerofoilportion of the blade radially outward of the root portion with thecooling channels branching off radially from the slit-like cooling airsupply channel into the blade profile of the aerofoil portion of theturbine blade.

An example for a cooling air inlet on the front side, i.e. the upstreamside, of the blade root is shown in EP 3 121 373 A1. Here, the coolingair inlet is designed with a geometry that forms a micro-compressor inorder to reduce pressure losses of the cooling air at the entrance intothe cooling air supply channel.

For efficient cooling of the blade it may also be desirable to have aninlet for cooling in the rear, i.e. on the axial downstream side, of theblade. However, experiments have shown undesired turbulence when thecooling air boards the spinning blade. The turbulence cause pressureloss, which in turn limits the amount of driving pressure available topush the coolant through the blade. This limits the extent of coolingpossible, which in turn limits the maximum operating temperature,thereby limiting both the life of the blade and the engine efficiency.

It is an object to form a cooled turbine rotor blade for a gas turbineengine, in particular an aircraft turbine engine, and a turbine rotor ofthe kind as it has been described above in such a manner that aneffective cooling of the turbine blades and of the disk rim and thushigh operating temperatures as well as a long service life of theturbine blades and a high engine efficiency can be ensured, when coolingair enters the spinning turbine rotor blade at its rear side.

The object is achieved with a cooled turbine rotor blade according tothe features of patent claim 1 and with a rotor-stator stage for a gasturbine according to the features of patent claim 10.

Further features, advantages and measures are listed in the sub-claims.The features and measures listed in the sub-claims can be combined withone another in advantageous ways.

According to an aspect, there is provided a cooled turbine rotor bladefor a gas turbine engine, in particular an aircraft turbine engine,comprising an aerofoil portion and a profiled root portion radiallyinward of the aerofoil portion in installed state, a platform betweenthe aerofoil portion and the root portion, and at least one cooling airsupply channel that extends through the platform into the interior ofthe aerofoil portion and therein up to at least one outlet opening,wherein an inlet opening of the cooling air supply channel is located atthe rear side of the rotor blade, and an inlet part of the cooling airsupply channel with the inlet opening is angled into the direction ofrotation (during operation) of the rotor blade and curved into aradially outward direction.

In use, the blade may have an upstream side and a downstream sidedefined relative to a flow direction through the gas turbine. Theupstream side may correspond to the leading edge of the blade anddefines the front side of the rotor blade, while the downstream side maycorrespond to the trailing edge of the blade and defines the rear sideof the rotor blade.

The air pressure at the rear is lower than the front due to the pressuredrop across the rotor stage. By using the low pressure side as thecooling air source, less compressor work is required to deliver thecooling air, so the overall engine efficiency is improved. Due toreduced compression, the cooling air is also cooler, so the bladeoperating temperature can be increased and/or less cooling flow can beused, both resulting in further engine efficiency improvements.

The arrangement of the inlet part of the cooling air supply channel onthe rear side of the rotor blade being angled into the rotationdirection of the rotor blade has the advantageous effect of a smoothcooling air boarding without additional pressure losses, and therefore,results in an efficient turbine blade cooling. Since the rotor blade isin operation spinning faster than the cooling air, the inlet part of thecooling air supply channel may be profiled to cause a rise of pressure,enabling use of more advanced aerofoil cooling techniques. In this way,the service life of the turbine blade and of the turbine rotor can beconsiderably increased while high operating temperatures can be used.

In an embodiment, the inlet part of the cooling air supply channel mayprovide a serpentine design that curves smoothly between the inlet andthe cooling air channel, and that may form at least substantially anS-shape. For example, the S-shape may have a first bend after boardingthe cooling air, and a second bend leading into a cooling channelradially outwards, with each bend having a smooth angle of more than90°. With such a design, the inlet part of the cooling air supplychannel smoothly turns the cooling air into the cooling air supplychannel with minimal pressure losses.

Further, the inlet part of the cooling air supply channel may comprise ageometry that forms a compressor, e.g. with widening diffuser-likediameter, in order to prevent pressure losses. In an expedientembodiment, the inlet of the cooling air may provide a channel geometrythat expands in flow direction in a diffuser-like manner. In this way, acompressor is created in a constructionally simple manner. Thedeceleration of the flow velocity that is taking place inside thecompressor leads to a steady increase in static pressure, whichultimately leads to a better cooling system.

Having the described low loss boarding of cooling air available opens upan opportunity to use a low pressure coolant source for high temperatureturbines, in particular the air cavity between a rotor stage of a gasturbine and a stator stage following downstream.

The platform may provide at the rear side of the rotor blade a rearwardprojecting overhang portion forming an air flow discourager, wherein theinlet opening of the cooling air supply channel is located radiallyinwards of the overhang portion.

The inlet opening position may be positioned at the platform radiallyoutwards of the root portion and optionally of an axial securing devicefor the root portion at the radial inward side of the overhang portionof the blade platform.

The rear flow discourager of a rotor blade typically forms sealingfeatures of a seal between a rotor stage of a gas turbine and afollowing stator stage which may combine or interact with neighbouringsurfaces for example of neighbouring vanes to form a seal whenassembled. The air cavity feeding the blade may be separated from theannulus rim by just the discourager, so the pressure is roughly the sameas in the annulus.

However, the disclosure is not limited to such a positioning of theinlet opening of the cooling air supply channel. The person skilled inthe art will make use of the advantages of the disclosure depending onthe application case, also with a higher or lower radial position of theinlet of the cooling air supply channel.

Further, the rotor blade may have a root portion that is shaped with afir-treelike profile, so that at least one additional axial cooling airchannel may be provided in the root portion or in a gap formed betweenthe root portion and a disk finger groove holding the root portion ofthe rotor blade on a turbine rotor disk. From such axial cooling airchannel one or more radial cooling passages may branch into the aerofoilportion of the turbine blade.

The aerofoil portion may include aerodynamic surfaces that, inoperation, are gas-washed by the working fluid, i.e. air, such as asuction surface, a pressure surface, and a platform from which thesuction surface and pressure surface extend. The rotor blade mayoptionally include a shroud at its tip. The shroud may include sealingfeatures to prevent or reduce over-tip leakage flow. The shroud mayextend around the entire outer circumference of the blades.Alternatively, the blade may have a partial shroud, or winglet, at itstip (for example extending around a portion of the segment between theblades).

To cause fluid to flow into the blade, the outlet opening of the coolingair supply channel is best to be located high up the blade e.g. at thetip of the aerofoil portion so the rotating cooling air supply channelhas a centrifugal pumping effect.

For most effective pumping of cooling air a radially straight upextending design for the cooling air channel may be provided. However,due to the minimized pressure loss at the entering of cooling air thereis enough pressure to push the cooling air also through a morecomplicated blade cooling system. Such a blade cooling system mayprovide an internal multipass passage with pedestals or turbulators, andthroughtrailing-edge ejection holes. Each pass may be arranged to carrycooling air in either a substantially radially outward direction or asubstantially radially inward direction. The outlet opening position maybe after at least one pass through the internal cooling passage from thecooling air inlet opening. For example, the outlet position may be aftertwo passes through the internal cooling passage, e.g. one pass in theradially outward direction and one pass in the radially inwarddirection. The radially outward and radially inward directions may bealigned with a longitudinal direction, or chordwise direction, of theblade or the aerofoil portion thereof. The radially outward directionmay correspond to a root-to-tip direction of the aerofoil. The radiallyoutward and radially inward directions may refer to the directionsrelative to a gas turbine engine (for example an axial flow gas turbineengine), for example when the aerofoil is installed in the gas turbineengine.

Again, this augments the extent of cooling possible, which in turnincreases the maximum operating temperature, thereby increasing both thelife of the blade and the engine efficiency.

According to an aspect, there is provided a rotor-stator stage for a gasturbine, wherein the rotor stage comprises at least one rotor bladeaccording to any one of the preceding claims and an air cavity radiallyinwards of sealing features between the rotor stage and a neighbouringdownstream stator stage forms the source for feeding the rotor bladewith cooling air.

Such a rotor-stator stage may provide any one or more of the advantagesdescribed herein.

The suggested turbine rotor blade as well as the suggested rotor-statorstage are not necessarily limited to aero engines, their benefits can berealised on any gas turbine e.g. in marine or energy applications.

Apart from the mentioned combination of features, the features specifiedin the patent claims as well as the features specified in the followingexemplary embodiment are suitable respectively on their own or in anycombination with each other to embody the subject matter according tothe disclosure.

Other advantages and embodiment possibilities of the cooled turbineblade according to the disclosure also follow from the patent claims andthe exemplary embodiment that will be described in principle below byreferring to the drawing, in which:

FIG. 1 shows a portional side view of a gas turbine engine providingcooled turbine rotor blades and a rotor-stator stage according to thedisclosure;

FIG. 2 shows a partial portion view of a rotor-stator stage with aturbine rotor blade providing an internal cooling air supply channel;

FIG. 3 shows a simplified perspective view the rotor blade of FIG. 2seen in the direction of arrow P in FIG. 2; and

FIG. 4 shows a perspective view of a part of the rotor blade of FIG. 2and FIG. 3 with an inlet opening of a cooling air supply channel.

With reference to FIG. 1, a ducted fan gas turbine engine 10, here foruse in an aircraft, has a principal and rotational axis X-X. The engine10 comprises, in axial flow series, an air intake 11, a propulsive fan12, an intermediate pressure compressor 13, a high-pressure compressor14, combustion equipment 15, a high-pressure turbine 16, an intermediatepressure turbine 17, a low-pressure turbine 18 and a core engine exhaustnozzle 19. A nacelle 21 generally surrounds the engine 10 and definesthe intake 12, a bypass duct 22 and a bypass exhaust nozzle 23.

The gas turbine engine 10 works in a conventional manner so that airentering the intake 11 is accelerated by the fan 12 to produce two airflows: a first air flow A into the intermediate pressure compressor 13and a second air flow B which passes through the bypass duct 22 toprovide propulsive thrust. The intermediate pressure compressor 13compresses the air flow A directed into it before delivering that air tothe high pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 14 isdirected into the combustion equipment 15 where it is mixed with fueland the mixture combusted. The resultant hot combustion products thenexpand through, and thereby drive, the high, intermediate andlow-pressure turbines 16, 17, 18 before being exhausted through thenozzle 19 to provide additional propulsive thrust. The high,intermediate and low-pressure turbines 16, 17, 18 respectively drive thehigh and intermediate pressure compressors 14, 13 and the fan 12 bysuitable interconnecting shafts.

As the air passes through the gas turbine engine 10 it is heated to hightemperatures. In particular, the first airflow B reaches hightemperatures as it passes, through the core of the engine. Typically,particularly high temperatures may be reached at the exit of thecombustion equipment 15, and as the air subsequently passes through thehigh, intermediate and low-pressure turbines 16, 17, 18.

Gas temperatures in the turbine can be in excess of 2100 K. It isdesirable to operate the turbine at the highest possible temperaturebecause generally, for a given gas turbine configuration, increasing theturbine entry gas temperature will produce more specific thrust.

Therefore, in order to cool turbine rotor blades 25, internal coolingair supply channels 26 may be formed within the blades 25. Theseinternal cooling air supply channels 26 allow cooling air to be passedthrough the blades to remove heat through convection.

Typically, cooling air 100 may be bled from the compressor 13, 14, priorto combustion, for example from the HP compressor. Typical cooling airtemperatures are between 700 and 900 K.

In addition to the cooling air 100 that is provided to the turbine rotorblade 25 from the compressor 13, 14 in the arrangement shown in FIG. 2to FIG. 4, more air is required for sealing purposes between a rotorstage 20 and a for example downstream neighbouring stator stage 21.

The turbine rotor blade 25 may be any type of turbine blade. Forexample, the turbine blade 25 may be part of a high pressure turbine 16,an intermediate pressure turbine 17, or a low pressure turbine 18. Theturbine blade 25 may be part of any type of gas turbine engine, forexample a ducted fan gas turbine (turbofan) engine 10 such as that shownin FIG. 1, a turbojet, a turboprop, a turboshaft, an open rotor engine,or any other gas turbine engine, for example axial flow or radial flow.

The turbine blade 25 has a root portion 30, a platform 31, an aerofoilportion 32, and a tip 33. Cooling air 100 passes through a cooling airsupply channel 26 from an inlet opening 40 through the platform 31 intothe interior of the aerofoil portion 32 and therein radially outwards upto an outlet opening 41 at the tip 33 of the rotor blade 25.

The root portion 30 may allow the rotor blade 25 to be attached to acorresponding component of a gas turbine engine, for example to aturbine disc 34. The term root portion 30 may be used to refer to partsof the rotor blade 25 that are radially inward of the platform 31. Theroot portion 30 has an attachment profile 35 which may be a fir tree asshown in FIG. 4. The attachment profile 35 allows the rotor blade 25 tobe attached to the corresponding turbine disc 34.

The rotor blade 25 may be said to be shrouded. However, different typesof blades may be used, for example also partially shrouded turbine rotorblades.

The inlet opening 40 of the cooling air supply channel 26 is located atthe rear side 36 of the rotor blade 25 at the platform 31. The platform31 as shown comprises at the rear side 36 of the rotor blade 25 arearward projecting overhang portion 37 forming an air flow discourager.The rearward overhang portion 37 forms sealing features of a seal 39between the rotor stage 20 and the following stator stage sealing an aircavity 27 from which the cooling air supply channel 26 is feeded.

In the shown embodiment, the inlet opening 40 of the cooling air supplychannel 26 is positioned at the beginning of the overhang portionradially inwards of the overhang portion 37 and radially outwards of theroot portion 30 and its axial securing device 38.

The inlet opening 40 and a following inlet part 42 of the cooling airsupply channel 26 are angled into the direction of rotation (inoperation) of the rotor blade, and curved into an radially outwarddirection. Thus, the inlet part 42 of the cooling air supply channel 26may provide a serpentine design that may form at least substantially aS-shape with a radially straight upward ending. Also more bends than inS-shape and/or sharper turns, e.g. 180° bends, may be chosen by theperson skilled in the art.

With such design, a smooth entry of cooling air 100 into the inletopening 40 and a pressure recovery within the inlet part 42 of thecooling air supply channel 26 is achieved, providing the afore-mentionedadvantages.

LIST OF REFERENCE SIGNS

-   10 Gas-turbine engine-   11 Air intake-   12 Fan-   13 Intermediate-pressure compressor-   14 High pressure compressor-   15 Combustion equipment-   16 High pressure turbine-   17 Intermediate pressure turbine-   18 Low pressure turbine-   19 Exhaust nozzle-   20 Rotor stage-   21 Stator stage-   22 Bypass duct-   23 Bypass exhaust nozzle-   25 Turbine rotor blade-   26 Cooling air supply channel-   30 Root portion-   31 Platform-   32 Aerofoil portion-   33 Tip-   34 Turbine disc-   35 Attachment profile, fir tree profile-   36 Rear side-   37 Overhang portion-   38 Axial securing device-   39 Seal-   40 Inlet opening-   41 Outlet opening-   42 Inlet part-   100 Cooling air-   A First air flow-   B Second air flow-   P View arrow-   X Axis

1. A cooled turbine rotor blade for a gas turbine engine, in particularan aircraft turbine engine, comprising an aerofoil portion and aprofiled root portion radially inward of the aerofoil portion ininstalled state, a platform between the aerofoil portion and the rootportion, and at least one cooling air supply channel that extendsthrough the platform into the interior of the aerofoil portion andtherein up to at least one outlet opening, wherein an inlet opening ofthe cooling air supply channel is located at the rear side of the rotorblade, and an inlet part of the cooling air supply channel with theinlet opening is angled into the direction of rotation of the rotorblade and curved into a radially outward direction.
 2. The cooledturbine rotor blade as claimed in claim 1, wherein the inlet part of thecooling air supply channel provides a serpentine design.
 3. The cooledturbine rotor blade as claimed in claim 1, wherein the inlet part of thecooling air supply channel forms at least substantially a S-shape. 4.The cooled turbine rotor blade as claimed in claim 1, wherein the inletpart of the cooling air supply channel comprises a geometry that forms acompressor diffuser.
 5. The cooled turbine rotor blade as claimed inclaim 1, wherein the inlet opening of the cooling air supply channel islocated at the platform at its radially inwards side.
 6. The cooledturbine rotor blade as claimed in claim 1, wherein the platform providesat the rear side of the rotor blade a rearward projecting overhangportion, wherein the inlet opening of the cooling air supply channel islocated radially inwards of the overhang portion.
 7. The cooled turbinerotor blade as claimed in claim 1, wherein at least one outlet openingof the cooling air supply channel is located at a tip of the aerofoilportion.
 8. The cooled turbine rotor blade as claimed in claim 1,wherein at least one cooling air supply channel is extending radiallystraight up in the aerofoil portion.
 9. The cooled turbine rotor bladeas claimed in claim 1, wherein the cooling air supply channel isconfigured as a multipass passage system.
 10. A rotor-stator stage for agas turbine, wherein the rotor stage comprises at least one rotor bladeaccording to claim 1, and an air cavity radially inwards of sealingfeatures between the rotor stage and a neighbouring downstream statorstage forms a source for feeding the rotor blade with cooling air.